Gas turbine engines generally include a turbine section downstream of a combustion section that is rotatable with a compressor section to rotate and operate the gas turbine engine to generate power, such as propulsive thrust. General gas turbine engine design criteria often include conflicting criteria that must be balanced or compromised, including increasing fuel efficiency, operational efficiency, and/or power output while maintaining or reducing weight, part count, and/or packaging (i.e. axial and/or radial dimensions of the engine).
Conventional gas turbine engines generally include turbine sections defining a high pressure turbine in serial flow arrangement with an intermediate pressure turbine and/or low pressure turbine. Additionally, conventional gas turbine engine turbine sections generally include successive rows or stages of stationary and rotating airfoils (e.g. vanes and blades). Stationary airfoils or vanes are often employed to direct or otherwise condition a flow of combustion gases before passing across rotating airfoils or blades. Stationary airfoils often require cooling air routed from other areas of the gas turbine engine, such as the compressor section, to mitigate damage from combustion gases. However, routing air from the compressor section to the turbine section, thereby bypassing the combustion section, generally removes energy for combustion and therefore reduces gas turbine engine efficiency.
Furthermore, conventional low pressure turbines often require a plurality of stages to distribute energy or work to operate the fan assembly and/or compressor to which the low pressure turbine is driving. However, the plurality of stages contribute to axial and radial dimensions of the gas turbine engine, which thereby contributes to weight of the overall engine and aircraft to which it is attached, and consequently adversely impacts fuel efficiency, engine performance, and engine and aircraft efficiency.
Known solutions include adding a reduction gearbox between a fan assembly and an engine core, which may reduce the quantity of the plurality of stages of a turbine section necessary to operate the fan assembly and compressor to which it is attached, and may generally provide some net increase in engine efficiency and improvement in fuel consumption. However, adding a reduction gearbox introduces new complexities and limitations to turbine engine design and operation. For example, known reduction gearboxes have an approximately 100% amount of torque or power routed in series from a low pressure turbine through the gearbox to drive a fan assembly. In such known arrangements, routing an approximately entire amount of torque or power from the low pressure turbine through the gearbox to the fan assembly necessitates complex gearbox designs, increased gearbox weight for the stresses and loads from the substantially entire load from the turbine section, and generally larger diameters of gearbox, thereby retaining or increasing radial dimensions of the engine.
Still further, known solutions including reduction gearboxes in which approximately 100% of torque or power from the low pressure turbine is directed through the gearbox to the fan assembly renders a systemic failure of the gearbox as a single point of failure. In such an arrangement, loss of gearbox operation results in loss of substantially all power from the low pressure turbine being delivered to the fan assembly, thereby reducing engine thrust or power output to an amount produced solely by the engine core through the core flowpath (e.g. approximately 10% of total thrust).
Therefore, there exists a need for an engine that may incorporate a reduction gearbox while reducing or eliminating adverse effects of gearbox placement, such as increased turbine engine packaging, such as increased diameter, axial length, or both, and/or single-point system failure of low pressure turbine power to the fan assembly.